The present invention relates generally to gas turbine engines, and, more specifically, to turbine nozzles therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel and ignited in a combustor for generating hot combustion gases. Energy is extracted from the gases in corresponding turbine stages which power the compressor and produce useful work, such as powering a fan in a turbofan engine for propelling an aircraft in flight, for example.
Since the turbines are bathed in the hot combustion gases during operation, they must be suitably cooled which is typically accomplished by bleeding a portion of the pressurized air from the compressor and channeling it through the turbine components.
A high pressure turbine directly receives gases from the combustor and includes a stator nozzle and a corresponding first stage rotor having a plurality of rotor blades extending radially outwardly from a supporting disk. A second stage nozzle then directs the combustion gases through a corresponding row of rotor blades extending from another rotor disk. The second stage nozzle receives lower temperature combustion gases than the first stage nozzle and therefore has different cooling requirements, which are typically effected in a different manner than that for the first stage nozzle.
Turbine nozzles are designed for durability with extensive lives measured in thousands of hours or thousands of cycles of operation. Such extended life is difficult to achieve since the nozzles are subject to various differential temperatures during operation which create thermal loads and stress therefrom. And, temperature distributions and heat transfer coefficients of the combustion gases channeled through the nozzle vary significantly and increase the complexity of providing corresponding cooling. Suitable nozzle cooling is required to limit thermal stresses and ensure a useful life.
A typical turbine nozzle includes a row of stator vanes joined at radially opposite ends to corresponding outer and inner bands. The bands are typically segmented in the circumferential direction, and include two or more vanes in corresponding sectors. The vane sectors permit differential movement during combustion gas temperature changes for reducing undesirable thermal stress during operation.
The individual vanes are hollow and typically include an impingement baffle therein which is a perforated sheet metal sleeve spaced from the inner surface of the vane cavity for channeling cooling air in impingement jets there against.
This type of turbine nozzle specifically configured for a second stage turbine has enjoyed many years of commercial service in this country. However, these nozzles are beginning to experience distress at high cycle operation which may require their replacement prior to their expected useful life. Nozzle distress is caused by locally high heat transfer coefficients in different regions of the nozzle at which corresponding cooling is limited. Thermal gradients lead to thermal stress, which adversely affect the useful life of the nozzle.
Accordingly, it is desired to uncover the source of high cycle turbine nozzle distress, and improve the nozzle design for increasing nozzle durability and corresponding life.